WO2006033407A1

WO2006033407A1 - Wall shape of axial flow machine and gas turbine engine

Info

Publication number: WO2006033407A1
Application number: PCT/JP2005/017515
Authority: WO
Inventors: Mizuho Aotsuka; Hiroshi Hamazaki; Akira Takahashi; Haruyuki Tanimitsu
Original assignee: Ishikawajima-Harima Heavy Industries Co., Ltd.
Priority date: 2004-09-24
Filing date: 2005-09-22
Publication date: 2006-03-30
Also published as: EP1760257A1; JPWO2006033407A1; EP1760257B1; EP1760257A4; US20070258810A1; CA2569026C; CA2569026A1; JP4640339B2; US7690890B2

Abstract

A wall shape of an axial flow machine capable of reducing a loss by a secondary flow, wherein a groove (40) extending in the axial direction of a cascade (30) is formed in the areas of a wall (31) between the blades (14) of the cascade (30). The area where the groove (40) is formed is positioned between the leading edges (20) and the trailing edges (21) of the blades (14) in the axial direction, and the centerline shape of the groove (40) comprises warpage in the same direction as that of the camber line (22) of the blades (40). The deepest part of the groove (40) is positioned between the near center parts or the center parts of the blades (14) and the leading edges (21) in the axial direction.

Description

Specification

Wall shape of axial flow machine and gas turbine engine

Technical field

[0001] The present invention relates to a shape of a wall facing a flow path of an axial flow machine.

This application claims priority based on Japanese Patent Application No. 2004-277114 filed on September 24, 2004, the contents of which are incorporated herein by reference.

Background art

In an axial flow machine having cascades such as a gas turbine engine, a flow path is sandwiched between radially inner and outer walls, and a boundary layer develops on the wall surface. In the wall boundary layer, a secondary flow with a velocity component different from the main flow occurs due to the pressure gradient between the blades. This secondary flow is known to cause pressure loss (energy loss).

[0003] For the purpose of reducing secondary flow loss, there is a technique for reducing a pressure gradient between blades by providing a gradient in a radial wall facing a flow path (see, for example, Patent Document 1).

Patent Document 1: US Pat.

Disclosure of the invention

Problems to be solved by the invention

[0004] In recent years, in an axial flow machine such as a gas turbine engine, the thickness of a blade tends to be reduced for the purpose of reducing the weight. In blade cascades using thin blades, if a separation zone is formed in the vicinity of the blade's abdominal surface (pressure surface; pressure surfa ce) immediately and interferes with the wall boundary layer, it has a strong axis with a flow direction axis different from the mainstream. A vortex is generated, causing pressure loss (energy loss). With conventional technology, this strength and loss due to vortices are not sufficiently reduced.

[0005] An object of the present invention is to provide a wall shape of an axial flow machine capable of reducing a loss due to a secondary flow, and a gas turbine engine.

Means for solving the problem

[0006] The wall shape of the axial flow machine of the present invention is a shape of a radial wall facing the flow path of the axial flow machine having a blade row, and is a region between the blades in the blade row, A groove extending in the axial direction of the blade row, and the groove formation region has a leading edge and a trailing edge of the blade with respect to the axial direction. The groove center line shape has a warp in the same direction as the wing warp line, and the deepest part of the groove is near the center of the wing or the center of the wing with respect to the axial direction. It is located between the front edge.

[0007] In the wall shape of the present invention, the deepest portion of the groove may be located 20 to 60% of the distance between the leading edge and the trailing edge of the blade from the leading edge of the blade with respect to the axial direction. preferable.

[0008] Further, it is preferable that the deepest portion of the groove is located at 30 to 50% of the distance between the front edge and the rear edge of the blade with respect to the axial direction.

In the wall shape of the present invention, it is preferable that at least a part of the center line of the groove is non-parallel to the warp line of the blade.

[0010] In the wall shape of the present invention, it is preferable that the force near the center of the blade is directed toward the vicinity of the trailing edge so that the groove force approaches the back surface of the blade.

[0011] In the wall shape of the present invention, it is preferable that the distance between the center line of the groove and the back surface of the blade is the shortest in the vicinity of the trailing edge of the blade.

In the wall shape of the present invention, it is preferable that the peripheral shape of the wall at the front edge position and the rear edge position of the blade is an arc.

In the wall shape of the present invention, the circumferential contour of the wall in the vicinity of the leading edge of the wing is

It is preferable to include a convex shape (positive curvature) adjacent to the abdominal surface of the wing and a convex shape (positive curvature) adjacent to the back surface of the wing! /.

[0014] In the wall shape of the present invention, the circumferential contour of the wall in the vicinity of the trailing edge of the wing is

Preferably, it includes a concave shape (negative curvature) adjacent to the back surface of the wing.

[0015] In this case, it is preferable that the circumferential contour of the wall in the vicinity of the trailing edge of the wing further includes a convex shape (positive curvature) adjacent to the abdominal surface of the wing.

In the wall shape of the present invention, it is preferable that the contour of the wall along the abdominal surface of the wing includes a convex region in the vicinity of the leading edge of the wing and a convex region in the vicinity of the trailing edge of the wing. Yes.

In the wall shape of the present invention, it is preferable that the contour of the wall along the rear surface of the wing includes a convex region near the front edge of the wing and a concave region near the rear edge of the wing. Yes. [0018] In this case, it is preferable that the concave region in the vicinity of the trailing edge of the wing is 50% or less of the chord length of the wing.

[0019] The gas turbine engine of the present invention is a gas turbine engine having a plurality of stationary blades and a plurality of moving blades, wherein a wall on a root side of the plurality of stationary blades and a tip side of the plurality of stationary blades are provided. At least one of the wall, the wall on the root side of the plurality of moving blades, and the wall on the tip side of the plurality of moving blades has the wall shape of the present invention.

[0020] Examples of the gas turbine engine include a turbofan engine, a turbojet engine, a turboprop engine, a turboshaft engine, a turbo ramjet engine, a gas turbine for power generation, and a marine gas turbine.

The invention's effect

[0021] According to the wall shape of the fluid machine of the present invention, the loss due to the secondary flow can be reduced. Further, according to the gas turbine engine of the present invention, the performance is improved by reducing the loss due to the secondary flow. Is planned.

Brief Description of Drawings

FIG. 1 is a schematic cross-sectional view showing a gas turbine engine used in an aircraft or the like as an example of an axial flow machine to which the present invention is applied.

FIG. 2 is a view showing an embodiment in which the wall shape of the present invention is applied to the wall on the root side of the rotor, and the surface height of the region between the blades is shown using contour lines.

FIG. 3 is a perspective view showing the vicinity of a wall of a blade row.

FIG. 4 is a diagram showing the depth shape of the groove.

FIG. 5A is an explanatory diagram of the definition of groove depth.

FIG. 5B is an explanatory diagram of another definition of groove depth.

FIG. 6A is a diagram for explaining the shape of a groove.

FIG. 6B is a diagram showing the circumferential shape of the wall in region A in FIG. 6A.

FIG. 6C is a diagram showing the circumferential shape of the wall in region C in FIG. 6A.

FIG. 7 is an explanatory diagram of the definition of the wing chord length.

[Figure 8] As a comparative example, the flow field near the wall when the blade wall is flat (the pine near the wall) It is a figure which shows (Ha number distribution).

FIG. 9A is a diagram showing a flow field near the wall (Mach number distribution near the wall surface) when the blade row wall is flat as a comparative example.

FIG. 9B is a diagram showing a flow field (wall surface vicinity Mach number distribution) in the vicinity of the wall in the wall-shaped example of the present invention.

FIG. 10 is a graph showing a loss due to a secondary flow.

Explanation of symbols

[0023] 10, 11, 12, 13 ··· Rotor (base), 14 ··· Wings (moving blades), 15, 16, 17, 18 ··· Casing

(Base), 19 ··· Wings (static vanes), 20 · · · Leading edge, 21 · · · Rear edge, 22 · · · Wing warp, 23 · · · Abdominal surface (PS), 24 · · · Back ( SS), 29 ... string, 30 ... cascade, 31 ... wall, 40 ... groove, 41 ... groove center line, 45 ... peeling zone, 46 ... vortex, 50, 51, 54 ··· convex shape, 55 ··· concave shape, 60, 61, 64, 65 ··· convex region, 66 ··· concave region.

BEST MODE FOR CARRYING OUT THE INVENTION

FIG. 1 is a schematic cross-sectional view showing a gas turbine engine (turbofan engine) used in an aircraft or the like as an example of an axial flow machine to which the present invention is applied.

The gas turbine engine includes an air intake 1, a fan 'low pressure compressor 2, a fan air discharge duct 3, a high pressure compressor 4, a combustion chamber 5, a high pressure turbine 6, a low pressure turbine 7, and an exhaust duct 8.

[0025] A plurality of fans' low pressure compressor 2, high pressure compressor 4, high pressure turbine 6 and low pressure turbine 7 are arranged on the outer peripheral surface of each of the rotors 10, 11, 12, 13 as a base. Rotor blades 14 are spaced apart from each other in the circumferential direction, and a plurality of blades (stator blades) on the inner peripheral surface of annular casings 15, 16, 17, 18 as the base 19 includes nozzles (nozzles, stators) that are spaced apart from each other in the circumferential direction.

A plurality of wings 14 extend outward from each rotor 10, 11, 12, 13 and a plurality of wings 19 ί and each casing 15, 16, 17, 18 force extend inward. /! An annular flow path (axial flow path) is formed between the rotors 10, 11, 12, 13 and the corresponding casings 15, 16, 17, 18, respectively.

[0026] In the fan 'low pressure compressor 2 and high pressure compressor 4, the flow of the working fluid in the axial flow path Along the pressure of the working fluid increases. In the high pressure turbine 6 and the low pressure turbine 7, the pressure of the working fluid decreases along the flow of the working fluid in the axial flow path.

[0027] The peripheral wall on the base (node, hub) side of the rotor blade 14 is a radially inner wall in the axial flow path.

The inner wall of the nozzle blade 19 on the base side of the nozzle blade 19 is a radially outer wall in the axial flow path.

When a peripheral wall (such as a shroud wall) is provided on the tip (tip) side of the rotor blade 14, the tip side wall is an outer end wall in the axial flow path. When a peripheral wall is provided on the tip (tip) side of the nozzle blade 19, the tip side wall is an inner end wall in the axial flow path.

The wall shape of the present invention may be any one of the wall on the root side of the rotor blade 14, the wall on the tip side of the rotor blade 14, the wall on the root side of the nozzle blade 19, and the wall on the tip side of the nozzle blade 19. It is also applicable to.

FIG. 2 is a view showing an embodiment in which the wall shape of the present invention is applied to the wall on the root side of the rotor, and the surface height (radial position, contour) of the region between the blades is shown using contour lines. It shows. FIG. 3 is a perspective view showing the vicinity of the blade row wall.

[0029] As shown in FIG. 2, each wing 14 has a leading edge 20, a trailing edge 21, a vent surface (pressure surface (PS)) 23, and a back surface (negative). Suction surface (SS)) 2 4, and blade row 30 has a warp (warp line 22) protruding in the same circumferential direction. Due to the warp line 22 of the wing 14, the axial flow cross section decreases the force near the center of the wing 14 toward the trailing edge 21 of the wing 14.

As shown in FIG. 2 and FIG. 3, grooves 40 are formed in regions between the blades 14 in the radial wall 31 of the blade row 30. The groove 40 extends at least in the axial direction (X direction) of the cascade 30. The formation region of the groove 40 is between the leading edge 20 and the trailing edge 21 of the blade 14 with respect to the axial direction. That is, the formation region of the groove 40 is within the length of the chord 29 of the wing 14. With respect to the axial direction, one end of the groove 40 is located near the leading edge 20 of the wing 14 and the other end is located near the trailing edge 21 of the wing 14.

Further, as shown in FIG. 2, the groove 40 is formed to be curved along the warp line 22 of the blade 14 as a whole. That is, the shape of the center line 41 of the groove 40 is the same direction as the warp line 22 of the wing 14. It has a warp (a warp protruding in the same circumferential direction of the blade row 30). At least a portion of the center line 41 of the groove 40 is non-parallel to the warp line 22 of the wing 14. In other words, the phase force of the shape of the groove 40 changes with respect to the chord direction of the blade 14.

More specifically, the groove 40 has a shape in which the force near the center of the blade 14 gradually approaches the back surface 24 of the blade 14 toward the vicinity of the trailing edge 21. Near the trailing edge 21 of the wing 14 (near the end of the chord 29), the distance between the center line 41 of the groove 40 and the back surface 24 of the wing 14 is the shortest. The shortest distance between the center line 41 of the groove 40 and the back surface 24 of the blade 14 is preferably 50% or less of the longest distance. Also, in the vicinity of the trailing edge 21 of the blade 14, the angle between the center line 41 of the groove 40 and the axis of the blade row 30 | 8 (the center line 41 of the groove 40 with respect to the axis of the blade row 30 at the outlet of the flow Is the angle between the warp line 22 of the blade 14 and the axis of the blade row 30 (the outlet angle) of the warp line 22 of the blade 14 relative to the axis of the blade row 30 at the outlet of the flow. Larger than the angle formed by the tangential direction (exit angle). In the vicinity of the leading edge 20 of the blade 14, the positional relationship between the warp line 22 of the blade 14 and the center line 41 of the groove 40 varies depending on the airfoil shape and the flow field. For example, the warp line 22 of the wing 14 and the center line 41 of the groove 40 may be formed so as to intersect (ie, have the shortest distance force ^) in the range of the leading edge 20 and the trailing edge 21! / ,.

FIG. 4 is a view showing the depth shape of the groove 40 (cross-sectional shape of the groove projected onto the plane including the X axis) along the axial direction (X direction) of the blade row 30.

As shown in FIGS. 4 and 2, the depth shape of the groove 40 is the deepest portion 43 (see FIG. 2) and the shallowest portions 44a and 44b (see FIG. 2) along the axial direction (X direction) of the blade row 30. Gradually change between The deepest portion 43 of the groove 40 is located near the center of the blade 14 or between the center of the blade 14 and the leading edge 20 in the axial direction. The shallowest portions 44a and 44b of the groove 40 are located in the vicinity of the leading edge 20 and the trailing edge 21 of the blade 14 in the axial direction.

[0034] Here, the distance between the leading edge 20 and the trailing edge 21 of the blade 14 (measured from the leading edge 20 along the axis of the blade row 30) is defined as the "axial chord length". .

The depth of the groove 40 is defined as the radial distance (TD1) of the axial flow path reference plane (cylinder base (cylindrical or conical)) force, as shown in FIG. 5A. Alternatively, the depth of groove 40 is defined as half 0 f peak to peak (HR1 shown in FIG. 5B) in one section perpendicular to the axis of cascade 30 as shown in FIG. 5B. . The deepest part 43 of the groove 40 is TD1 or When HR1 is used, it is located 20 to 60%, preferably 20 to 50%, more preferably 30 to 50% of the axial length in the axial direction.

Further, as shown in FIGS. 4 and 2, the groove 40 gradually becomes shallower in the extending direction from the deepest portion 43 to the shallowest portions 44a and 44b at both ends. . That is, the groove 40 starts from the shallowest part 44a near the leading edge 20 of the wing 14 and deepens between the leading edge 21 of the wing 14 and near the center while increasing the depth (the deepest part 43). It ends at the shallowest part 44b near the trailing edge 21 of the wing 14 with decreasing height. The contour of the groove 40 from the deepest part 43 to the shallowest parts 44a, 44b is uniformly smooth or non-uniformly smooth. As described above, since the distance between the center line 41 of the groove 40 and the back surface 24 of the wing 14 is relatively short near the trailing edge 21 of the wing 14, the center depth of the groove 40 is near the center of the wing 14. In the vicinity of the trailing edge, the force is deeper at the part far from the rear face 24 of the wing 14 and shallower at the part near the rear face 24.

FIG. 6A is a diagram for explaining the shape of the groove 40. FIGS. 6B and 6C show the circumferential shape of the wall 31 having the groove 40 (circumferential contour, ie, the cross section of the wall (orthogonal to the axis). FIG. 3 is a diagram (concave distribution pattern of wall surfaces) showing a shape of a cross section).

As shown in FIG. 6A, at the front edge position (LE) and the rear edge position (TE) of the blade 14, the wall 31 has an annular shape, and its circumferential shape (circumferential contour) is an arc. That is, the circumferential shape of the wall 31 at the leading edge position (LE) and the trailing edge position (TE) has no recess due to the groove 40.

[0037] Here, between approximately 30% and 40% of the axial chord length of the wall 31 is referred to as region A, and approximately 60% to 90% of the axial length of the wall 31 is referred to as region C. About 40% -60% of the axial length of the wall 31 (ie, between region A and region C) is referred to as region B. The circumferential shape (concave convex) of the wall 31 is defined by region A and region C, and region B is a transition region that varies depending on the airfoil and flow field. Further, the ranges of the region A and the region C are appropriately changed depending on the place where the wall shape of the present invention is installed, the airfoil, and the flow field. For example, the range of Area C (approximately 60% —90%) is 60% —90%, 60% —80%, 70% —90%, 70% —80%, 80% —90%, 70%- It can be set to 85%, 75% —90%, 80% —95%.

As shown in FIG. 6B, in region A, the circumferential shape (circumferential contour) of the wall 31 is that the convex portion 50 adjacent to the ventral surface 23 (PS) of the wing 14 and the wing 14 Adjacent to rear 24 (SS) Another convex portion 51 and a concave portion formed between the two convex portions. The convex Z-concave Z convex shape in this area A is referred to as the “first shape”. The convex part has a positive curvature, and the concave part has a negative curvature.

As shown in FIG.6C, in region C, the circumferential shape of the wall 31 includes a convex portion 54 adjacent to the ventral surface 23 (PS) of the wing 14 and a concave portion 55 adjacent to the rear surface 24 (SS) of the wing 14. The transition from the convex portion 54 to the concave portion 55 is smooth. The abdominal convexity Z back concave shape in this region C is referred to as “second shape”.

[0039] As shown in FIG. 6A, in region B, the circumferential shape of wall 31 smoothly changes from the first shape to the second shape.

The region between the leading edge position (LE) and region A (ie, about 0% to 30% of the axial length) is a transition region and is referred to as region L. In the region L, the circumferential shape of the wall 31 smoothly changes to the first shape of the region A (convex Z concave Z convex) at the leading edge position (LE).

The region between region C and the trailing edge position (TE) (ie about 90-100% of the axial length) is also a transition region and is referred to as region T. In the region T, the circumferential shape of the wall 31 smoothly changes from the second shape in region C (abdominal surface convex Z back concave) to the arc at the trailing edge position (TE). As described above, the groove 40 has the deepest portion 43 between 20% and 60% of the axial length in any of the region L, the region A, and the region B.

[0040] In addition, the contour of the wall 31 along the ventral surface 23 of the wing 14 (the contour on the ventral side) is the reference for the axial flow path in all areas except the leading edge position (LE) and trailing edge position (TE). It is higher than the plane (cylinder base (cylindrical surface or conical surface)). As shown in FIG.6A, the contour on the ventral side is a convex region 60 having a positive curvature in the vicinity of the leading edge 20 of the wing 14 and a convex region having a positive curvature in the vicinity of the trailing edge 21 of the wing 14. 61. A region between the convex region 60 and the convex region 61 is a transition region, and in this transition region, the contour on the ventral surface side smoothly changes from the convex region 60 to the convex region 61. In this transition region, a recess having a negative curvature may be formed.

[0041] The contour of the wall 31 along the back surface 24 of the wing 14 (the contour on the back side) is in the main region, higher than the cylin der base of the axial flow path, and in the airfoil and flow field. And a partial region where the height of the reference surface force changes accordingly. The partial area is a reference, depending on the airfoil and flow field. It changes to a position higher than the surface, almost the same position, or a lower position. As shown in FIG.6A, the contour on the back side is a convex region 64 having a positive curvature near the leading edge 20 of the wing 14, a convex region 65 having a positive curvature near the center, and a vicinity of the trailing edge 21. And a concave region 66 having a negative curvature. The recessed area 66 in the rear profile is less than 50% of the chord length of the wing 14. As shown in Fig. 7, the chord length here is the distance (CL2) between the tip of the leading edge of the wing and the tip of the trailing edge (CL2), or a straight line perpendicular to the straight line that touches the leading and trailing edges. Defined as the distance (CL1) between two points touching the edge and the trailing edge.

FIG. 8 and FIG. 9A show, as a comparative example, a flow field near the wall when the blade wall is flat.

It is a figure which shows (wall surface vicinity Mach number distribution).

8 and 9A, on the wall surface, a separation zone 45 occurs partially near the ventral surface (PS) of the wing 14 (near the center in the chord direction), which interferes with the wall boundary layer and Strong vortices 46 with different flow direction axes are generated. The starting edge of the vortex 46 is the interference between the separation zone 45 and the wall boundary layer, relatively close to the wing's ventral surface, and with respect to the axial direction, near the center of wing 14 or between the center and leading edge of the wing Located in. The arrival position of the vortex 46 is on the back of the wing and near the trailing edge.

FIG. 9B is a diagram showing a flow field (wall surface vicinity Mach number distribution) in the vicinity of the wall in the wall-shaped embodiment of the present invention shown in FIGS. 2 to 6C.

As apparent from FIG. 9B, in the wall-shaped example of the present invention, the vortex is weakened and the flow disturbance on the wall surface is less than in the comparative example. As a result, in the embodiment, the flow loss (pressure loss, energy loss) can be reduced.

FIG. 10 is a graph showing a change in loss due to the secondary flow. In Fig. 10, the horizontal axis shows the span (the radial height of the blade row), and the vertical axis shows the flow loss (loss factor).

As is apparent from FIG. 10, the wall-shaped example of the present invention has a smaller flow loss than the comparative example. In particular, the loss reduction is remarkable near the wall surface indicated by S in the figure.

[0047] Further, the wall shape of the present invention is obtained by adding the root-side wall of the blade 14 of the other stage of the rotor, the tip-side wall of the rotor blade 14, the root-side wall of the nozzle blade 19, and the nozzle blade 19 The loss reduction similar to the above was confirmed analytically in all cases applied to the wall on the tip side. [0048] Thus, according to the wall shape of the present invention, the groove between the blades weakens the vortex caused by the interference between the separation region on the ventral side of the blade and the wall boundary layer, and reduces the flow loss due to the vortex. can do.

The centerline shape of the groove has a warp in the same direction as the wing warp line, thereby avoiding the generation of another loss vortex in the groove.

Since the deepest part of the groove is located near the center of the blade or between the center of the blade and the leading edge with respect to the axial direction of the blade row, the deepest part of the groove is located near the position where the vortex is generated. Therefore, the groove curvature change in the transverse section (cross section perpendicular to the axis) is relatively large near the vortex generation position.

In addition, the wall profile (circumferential contour and profile along the ventral surface) near the front edge of the wing has a convex shape near the leading edge of the wing. The change in the groove curvature in the cross section perpendicular to the axis is relatively large.

When the curvature change at the bottom of the groove in the cross section is large, the flow on the ventral side of the wing is accelerated. The acceleration of this flow suppresses the development of the boundary layer on the ventral side of the wing. Also, on the ventral side of the wing, the pressure decreases due to the acceleration of the flow, so the pressure difference between the flank side of the wing and the back side near the trailing edge of the wing decreases, resulting in a cross flow between the wings. Weaken.

Near the trailing edge of the wing, the wall contour (circumferential contour and contour along the abdominal surface) near the back of the wing has a concave shape, so the pressure near the vortex arrival point is high. As the pressure near the vortex reaches, the vortex becomes weaker.

[0049] In the wall shape of the present invention, the position and shape of the groove (groove 40 shown in Fig. 2) are optimized and designed according to the generation position of the vortex generated when the wall is flat and the traveling axis thereof. Is preferred. In this case, for example, the deepest part of the groove may be in the vicinity of the start end of the vortex. In addition, the extending direction of the groove in the vicinity of the trailing edge of the blade should be approximately approximate to the axial direction of the vortex. The wall shape shown in Fig. 2 to Fig. 6C is an example, and the wall shape of the cascade is appropriately optimized according to the airfoil shape and flow field.

[0050] While the preferred embodiments of the present invention have been described above, the present invention is not limited to these embodiments. Additions, omissions, substitutions, and other modifications can be made without departing from the spirit of the present invention. The present invention is not limited by the above description. Limited only by the scope of the appended claims.

Claims

The scope of the claims

[1] The shape of a radial wall facing the flow path of an axial flow machine having cascades,

A region extending between the blades in the blade row and extending in the axial direction of the blade row, wherein the groove formation region is between the leading edge and the trailing edge of the blade with respect to the axial direction. And the center line shape of the groove has a warp in the same direction as the warp line of the wing,

The deepest part of the groove is located near the center of the blade or between the center and the leading edge of the blade with respect to the axial direction.

[2] The deepest portion of the groove is located 20 to 60% of the distance between the leading edge and the trailing edge of the blade from the leading edge of the blade with respect to the axial direction. The wall shape of the axial flow machine described in 1.

[3] The deepest portion of the groove is located at 30 to 50% of the distance between the leading edge and the trailing edge of the blade with respect to the axial direction from the leading edge of the blade. The wall shape of the axial flow machine described in 1.

4. The wall shape of the axial flow machine according to claim 1, wherein at least a part of the center line of the groove is non-parallel to the warp line of the blade.

[5] The wall shape of the axial flow machine according to claim 1, wherein the force near the center of the blade is directed toward the vicinity of the trailing edge, and the groove approaches the back surface of the blade.

6. The wall shape of the axial flow machine according to claim 1, wherein the distance between the center line of the groove and the back surface of the blade is the shortest in the vicinity of the trailing edge of the blade.

7. The wall shape of the axial flow machine according to claim 1, wherein the circumferential shape of the wall at the front edge position and the rear edge position of the blade is an arc.

8. The circumferential contour of the wall in the vicinity of the leading edge of the wing includes a convex shape adjacent to the abdominal surface of the wing and a convex shape adjacent to the back surface of the wing. Wall shape of the axial flow machine described in.

9. The wall shape of the axial flow machine according to claim 1, wherein a circumferential contour of the wall in the vicinity of a trailing edge of the blade includes a concave shape adjacent to a back surface of the blade.

10. The wall shape of the axial flow machine according to claim 9, wherein the circumferential contour of the wall in the vicinity of the trailing edge of the wing further includes a convex shape adjacent to the abdominal surface of the wing.

[11] The contour of the wall along the abdominal surface of the wing includes a convex region near the leading edge of the wing and a convex region near the trailing edge of the wing. Wall shape of the described axial flow machine.

12. The contour of the wall along the back surface of the wing includes a convex region near the leading edge of the wing and a concave region near the trailing edge of the wing. Wall shape of the described axial flow machine.

13. The wall shape of the axial flow machine according to claim 12, wherein the concave region in the vicinity of the trailing edge of the wing is 50% or less of a chord length of the wing.

[14] A gas turbine engine having a plurality of stationary blades and a plurality of moving blades,

At least one of a plurality of stationary blades, a plurality of stationary blades, a plurality of stationary blades, a plurality of stationary blades, a plurality of stationary blades, a plurality of stationary blades, a plurality of stationary blades, and A gas turbine engine having the wall shape according to any one of claims 1 to 13.